EP1772742B1 - Correction of the distance between phase centres of two directional antenneas of a navigational satellite - Google Patents
Correction of the distance between phase centres of two directional antenneas of a navigational satellite Download PDFInfo
- Publication number
- EP1772742B1 EP1772742B1 EP06255096A EP06255096A EP1772742B1 EP 1772742 B1 EP1772742 B1 EP 1772742B1 EP 06255096 A EP06255096 A EP 06255096A EP 06255096 A EP06255096 A EP 06255096A EP 1772742 B1 EP1772742 B1 EP 1772742B1
- Authority
- EP
- European Patent Office
- Prior art keywords
- satellite
- navigational
- time period
- lever arm
- over
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Not-in-force
Links
Images
Classifications
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
- G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
- G01S19/01—Satellite radio beacon positioning systems transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
- G01S19/02—Details of the space or ground control segments
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
- G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
- G01S19/01—Satellite radio beacon positioning systems transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
- G01S19/03—Cooperating elements; Interaction or communication between different cooperating elements or between cooperating elements and receivers
- G01S19/07—Cooperating elements; Interaction or communication between different cooperating elements or between cooperating elements and receivers providing data for correcting measured positioning data, e.g. DGPS [differential GPS] or ionosphere corrections
-
- H—ELECTRICITY
- H01—ELECTRIC ELEMENTS
- H01Q—ANTENNAS, i.e. RADIO AERIALS
- H01Q1/00—Details of, or arrangements associated with, antennas
- H01Q1/12—Supports; Mounting means
- H01Q1/125—Means for positioning
-
- H—ELECTRICITY
- H01—ELECTRIC ELEMENTS
- H01Q—ANTENNAS, i.e. RADIO AERIALS
- H01Q3/00—Arrangements for changing or varying the orientation or the shape of the directional pattern of the waves radiated from an antenna or antenna system
- H01Q3/005—Arrangements for changing or varying the orientation or the shape of the directional pattern of the waves radiated from an antenna or antenna system using remotely controlled antenna positioning or scanning
Definitions
- the present invention relates generally to space-based navigational systems, and in particular, to a space-based lever arm correction system applicable to navigational systems employing spot beams.
- GPS Global Positioning System
- the GPS navigation signal is transmitted via a wide beam satellite antenna disposed on each GPS satellite.
- the wide beam antenna permits any GPS receiver having a line-of-sight to the GPS satellite to receive the navigational signal, and when the navigation signal from a sufficient number of GPS satellites has been acquired, the GPS receiver can determine its position via a precision clock and well-known triangulation techniques.
- a steerable high gain antenna may be used to transmit high intensity GPS signals via spot beams to areas where needed.
- the GPS system must provide users with real-time corrections of the spot beam antenna phase center location relative to the Earth coverage antenna phase center, even while the spot beam antenna is moving to track specific terrestrial locations.
- the present invention discloses a method and apparatus for estimating a lever arm correction between a wide beam antenna and a spot beam antenna.
- the method comprises the steps of receiving a target spot beam position over the time period T in the satellite; generating a predicted position of the satellite over a time period T; generating predicted satellite attitude over the time period T from a prescribed satellite attitude maneuver profile over the time period T; generating predicted gimbal angles of the spot beam antenna over the time period T from the predicted satellite attitude and the target spot beam position over the time period T; predicting the lever arm correction between the wide beam antenna and the spot beam antenna over the time period T from the predicted gimbal angles of the spot beam antenna over the time period T and the predicted satellite attitude over the time period T; and transmitting the predicted lever arm correction from the satellite to a navigational receiver receiving the second navigation signal.
- the apparatus can be embodied in a navigational satellite, which comprises a first antenna transmitting a first navigational signal; a steerable second antenna transmitting a second navigational signal; a satellite receiver, for receiving data from a ground station, the data including a target spot beam position over a time period T; a lever arm correction module, for predicting a lever arm correction between the first antenna and the second antenna, and a navigation signal transmitter for transmitting the predicted lever arm correction to a navigational receiver.
- FIG. 1A illustrates a three-axis stabilized satellite or spacecraft 100.
- the satellite 100 has a main body 102 (which may be referred to as the "satellite bus"), one or more solar panels 104, one or more navigation beam antennas 106E and 106S, and a telemetry and command antenna 108 which is used to communicate with a control ground station.
- the satellite 100 may also include one or more sensors 110 to measure the attitude of the satellite 100. These sensors may include sun sensors, earth sensors, and star sensors. Since the solar panels are often referred to by the designations "North” and “South”, the solar panels in FIG. 1 are referred to by the numerals 104N and 104S for the "North” and “South” solar panels, respectively.
- the three axes of the spacecraft 100 are shown in FIG. 1A .
- the pitch axis lies along the line of the solar panels' 104N and 104S mutual rotation axes.
- the roll and yaw axes are perpendicular to the pitch axis and lie in the directions and planes shown.
- the satellite 100 includes a first navigation beam antenna 106E and a second navigation beam antenna 106S.
- the first navigation beam antenna 106E is a wide-beam antenna which transmits a navigation signal with a beamwidth covering the widest range of the Earth's surface possible from that satellite's altitude at any time, and is directed toward the Earth along the yaw axis. Since this antenna 106E offers the widest coverage of the Earth's surface, it is typically not steerable.
- the navigation system uses a constellation of such satellites 100 to provide coverage anywhere on the Earth's surface by at least 4 different satellites at all times, thus permitting the navigational signal transmitted by the satellites to be used to determine the location of the receiver using triangulation techniques.
- the second navigation beam antenna 106S is a steerable spot beam antenna that provides a second navigation signal in a much narrower navigation beam. This allows transmission of a higher-strength beam to selected points on the ground without requiring excessive transmitter power, thus reducing the effectiveness of countermeasures such as jamming. Since the required service area includes substantially the entire surface of the Earth and the beamwidth of the spot beam antenna 106S is not wide enough to cover the entire surface area, the boresight of the spot beam antenna 106S can be steered about to direct the spot beam where desired. Such steering can be accomplished mechanically, by use of a gimbal structure driven by gimbal motors, or electronically, using phased arrays, for example.
- FIG. 1A also shows the phase center 112E of the wide beam antenna 106E and the phase center 112S of the spot beam antenna 106S. Since the spot beam antenna 106S is larger and offset from the wide beam antenna 106E, the phase centers 112S, 112E of the antennas are separated by an antenna lever arm 114, which is represented as a vector originating at the phase center 112E of the wide beam antenna 106E and extending to the phase center 112S of the spot beam antenna 106S. Due to motion of the satellite bus 102 and the spot beam antenna 106S and other factors, the antenna lever arm 114 does not remain fixed, but can vary substantially over time. This variance is enough to add a significant uncertainty in the ability of a navigation receiver (such as a GPS receiver) to determine its location when a navigation signal is received via the spot beam.
- a navigation receiver such as a GPS receiver
- the spot beam antenna 106S may be steered electronically (by appropriate phasing of elements in a scanning array) or mechanically (by use of a non-scanning antenna and a gimbal assembly), or a combination of both.
- FIG. 1B illustrates an embodiment using a gimbal assembly 152 having an inner gimbal 154A and an outer gimbal 154B.
- the inner gimbal 154A is associated with inner gimbal coordinate frame 156B while the outer gimbal 154B is associated with an outer gimbal coordinate frame 156C.
- the inner and outer gimbals 154A, 154B angularly direct the antenna 106S boresight (which is represented in antenna signal boresight coordinate frame 156D) where desired to transmit the navigation signal spot beam using the antenna 106S.
- Both inner and outer gimbals 154 typically include a potentiometer or other means to measure the gimbal angle.
- the spacecraft 100 determines its attitude via the attitude sensors 110, which may be mounted on the satellite bus or body 102. Using the measured satellite attitude and the angular and translational displacement between the satellite body 102 and the spot beam antenna 106S (as expressed by coordinate systems 156A-156D), the satellite 100 can determine the appropriate gimbal angles, and commands the gimbal motors to move the gimbals 156 to the appropriate positions.
- FIG. 2 is a diagram depicting the functional architecture of a representative attitude control system 200.
- Control of the spacecraft is provided by a computer or spacecraft control processor (SCP) 202.
- the SCP 202 performs a number of functions which may include post ejection sequencing, transfer orbit processing, acquisition control, station-keeping control, normal mode control, mechanisms control, fault protection, and spacecraft systems support, among others.
- the SCP 202 may implement one or more processing modules such as antenna control module 276, which is used to control the satellite spot beam antenna drive 274 to slew the spot beam antenna 106S to the appropriate orientation and to transmit a navigation signal.
- antenna control module 276 can be implemented in a different processor or in dedicated circuitry.
- Input to the spacecraft control processor 202 may come from any combination of a number of spacecraft components and subsystems, such as a transfer orbit sun sensor 204, an acquisition sun sensor 206, an inertial reference unit 208, a transfer orbit Earth sensor 210, an operational orbit Earth sensor 212, a normal mode wide angle sun sensor 214, a magnetometer 216, and one or more star sensors 218.
- the SCP 202 generates control signal commands 220 which are directed to a command decoder unit 222.
- the command decoder unit operates the load shedding and battery charging systems 224.
- the command decoder unit also sends signals to the magnetic torque control unit (MTCU) 226 and the torque coil 228.
- MTCU magnetic torque control unit
- Wheel torque commands 262 are generated by the SCP 202 and are communicated to the wheel drive electronics 240 which command the speed of the reaction wheels in reaction wheel assembly(s) 244.
- the spacecraft 100 includes four reaction wheels, at least one in each orthogonal direction, and one for redundancy purposes.
- the speed of the reaction wheels is also measured and fed back to the SCP 202 by feedback control signal 264.
- the SCP 202 also communicates commands and data 254 with command stations (further described in connection with FIG. 7 ) via a satellite transmitter/receiver (or transceiver) 258.
- the SCP 202 may include or have access to memory 270, such as a random access memory (RAM).
- memory 270 such as a random access memory (RAM).
- the SCP 202 operates under control of an operating system 272 stored in the memory 270, and interfaces with the other system components to accept inputs and generate outputs, including commands.
- Applications running in the SCP 202 access and manipulate data stored in the memory 270.
- the spacecraft 100 may also comprise an external communication device such as a satellite link for communicating with other computers at, for example, a ground station. If necessary, operation instructions for new applications can be uploaded from ground stations.
- instructions implementing the operating system 272, application programs, and other modules are tangibly embodied in a computer-readable medium, e.g., data storage device, which could include a RAM, EEPROM, or other memory device.
- the operating system 272 and the computer program are comprised of instructions which, when read and executed by the SCP 202, causes the spacecraft processor 202 to perform the steps necessary to implement and/or use the present invention.
- Computer program and/or operating instructions may also be tangibly embodied in memory 270 and/or data communications devices (e.g. other devices in the spacecraft 100 or on the ground), thereby making a computer program product or article of manufacture according to the invention.
- program storage device "article of manufacture” and “computer program product” as used herein are intended to encompass a computer program accessible from any computer readable device or media.
- FIG. 3 is a diagram depicting the geometrical relationships between the satellite 100, the Earth 308 and the Sun 306.
- Each satellite 100 of the GPS system is in an orbit 302 around the Earth 308, thus defining an orbital plane 304.
- the angle ⁇ between the orbital plane 304 and the Sun 306 is the Sun angle.
- the satellite 100 orbits the Earth 308, always directing the antenna 106E at the Earth 308.
- the satellite 100 attitude and solar panels 104 are adjusted to keep the solar panels 104 directed at the Sun 306 and the antenna 106E directed at the Earth 308 at the same time. Typically, this requires the satellite 100 to perform a large rotation about the yaw axis at the orbital midnight (0 degree orbit angle) and the orbital noon (180 degree orbit angle).
- the smaller the Sun angle ⁇ the larger the required rotation about the yaw axis.
- FIGs. 4A-4D are diagrams depicting the yaw turn maneuver at orbital midnight for a zero Sun angle (wherein the Sun 306 is in the orbital plane 304).
- FIG. 4A illustrates the satellite 100 attitude and solar panel 104 orientation as the orbital midnight is approached. Note that the solar panels 104 are canted at an angle to direct them to be perpendicular to the Sun's rays.
- the satellite 100 begins its maneuver and yaws counter clockwise as shown in FIG. 4B through FIG. 4D .
- the satellite 100 has yawed 180 degrees from its initial position.
- FIG. 5 is a diagram plotting the ideal yaw angle as a function of the orbit angle.
- the yaw angle movement is gradual, as shown in trace 502.
- the yaw angle abruptly transitions 180 degrees at orbital angles of 0 and 180 degrees, with resulting high yaw angular rates and/or accelerations, as shown in trace 504.
- Such low Sun angle conditions e.g. less than about five degrees
- FIG. 6 is a diagram illustrating an exemplary embodiment of a technique for estimating the lever arm 114 correction between the wide coverage antenna 106E and the spot beam antenna 106S.
- a prediction of the satellite 100 position over a time period T is generated.
- Block 604 illustrates the generation of a prediction of the satellite 100 attitude over the time period T from a profile of a prescribed navigational satellite attitude maneuver.
- block 606 illustrates predicting the lever arm correction between the wide coverage antenna and the spot beam antenna over the time period T using the predicted satellite attitude over the time period T and a target position on the Earth of the spot beam signal.
- FIG. 7 is a diagram illustrating architectural elements of a space-based navigational system 700.
- the navigational system 700 comprises a space segment 704 having a plurality of navigational satellites 100 (hereinafter alternatively referred to as spacecraft) in orbit around the Earth.
- Each of the satellites 100 includes a first transmitter 710, which transmits a first navigational signal using the wide beam antenna 106E and a second transmitter 712, which, using the spot beam antenna 106S, transmits a spot beam 719 having a second navigational signal to one or more of the navigational receivers 716 that together form the user segment 706.
- the spot beam 719 covers an area 715 on the Earth's surface, and is centered on the target spot beam position 717.
- Both the first navigational signal and the second navigational signal include information that can be used to determine the location of the navigational receiver 716.
- a navigation receiver may lose lock on the first navigation signal and will not be able to produce any positional solution.
- the second navigational signal is typically of greater power by virtue of being transmitted via a spot beam, and, will be used for positional determination using the navigational receiver 716 in the jamming environment.
- the satellites 100 in the space segment 704 are controlled via commands from the ground segment 702, which comprises one or more ground-based command stations 707, each including a command center 708 and a communicatively coupled antenna 714 for transmitting and receiving information and commands to and from the satellites 100 of the space segment 704.
- the operations illustrated in FIG. 6 can be performed on-board the satellite 100, the ground segment 702 controlling the satellite 100, by the navigational receivers 716 making up the user segment 706, or any combination thereof.
- the first approach that of transmitting the fixed antenna lever arm position in the satellite 100 coordinate frame in the navigational message and relying upon the GPS receivers to use this information along with a satellite yaw steering maneuver to enable computation of the corrections in a GPS coordinate frame
- the second approach makes predictions at least 15 minutes into the future, and during those 15 minutes, without additional information or constraints, this solution can yield large errors.
- FIGs. 8A and 8B are diagrams illustrating exemplary process steps that can be used to implement one embodiment of the invention, and how these process steps can be allocated to the architectural elements (e.g. the control segment 702, the space segment 704, and the user segment 706).
- the architectural elements e.g. the control segment 702, the space segment 704, and the user segment 706.
- the predicted position of the satellite 100 over a time period T is computed, as shown in block 802.
- the predicted position of the satellite 100 can be determined using a variety of methods well known in the art, and can include parameters to model satellite maneuvers (whether for mission specific or station keeping purposes). Also, for a navigational satellite such as a GPS satellite, the parameters for predicting satellite position are normally present on-board the satellite as an integral part of the navigational system. The predicted position of the satellite 100 can be determined from this information.
- a prediction of the satellite's attitude over the same time period T is computed. This is accomplished using the computed satellite position (from the computation shown in block 802), and from a prescribed satellite maneuver profile 808 that can be obtained either from the ground segment 702 or the memory of the satellite 100 in the space segment 704.
- FIG. 9 is a plot of the yaw attitude angle of the satellite 100 as a function of the orbit angle 310 and the Sun 306 angle for an exemplary prescribed satellite 100 maneuver profile.
- the prescribed maneuver profile can be defined such that, when accounting for the capabilities and limitations of the satellite's attitude control system 200, the satellite 100 can execute the prescribed attitude maneuver within a specific angular tolerance, the selection of such tolerance is a function of the desired lever arm estimation accuracy. In one embodiment, the value of this tolerance, expressed as the difference between the predicted satellite attitude and the actual satellite attitude to less than 0.2 degrees.
- the prescribed maneuver profile can be defined so as to limit the maximum angular velocity and/or acceleration about all axes (pitch, roll, and yaw) not to exceed maximum values or such that the maximum angular velocity and/or acceleration does not exceed maximum values in any one of the axes (for example, in the yaw axis).
- trace 502 illustrates yaw angle profile for a high Sun 306 angle ⁇ .
- the required yaw rates and accelerations are relatively small.
- the required yaw rates and accelerations are much higher, as shown in trace 504.
- These yaw rates are high enough so that the satellite's attitude control system 200 cannot follow the ideal profile, and thus, the deviation from the expected satellite 100 attitude angle and the actual satellite 100 attitude angle is significant enough to negatively impact the estimation of the lever arm 114.
- the maneuver is initiated based upon sensor (e.g. Sun, Earth, and/or star sensors) measurements, hence the time the maneuver is initiated is another source of unpredictability. To prevent this, the satellite is commanded to perform a prescribed satellite maneuver 808.
- This prescribed satellite maneuver profile 808 can be defined such that, when accounting for the capabilities and limitation of the satellite's attitude control system 200, the satellite 100 can execute the prescribed attitude maneuver to within specified tolerances.
- the maneuver profile can be arrived at by (1) limiting the satellite attitude angular rate and/or acceleration in all axes (e.g. pitch, roll, and yaw) or a single axis (e.g. yaw only) to a maximum value, (2) limiting the error between the predicted satellite attitude and the actual satellite attitude (when attitude control system 200 limitations are considered), (3) and/or placing other suitable restrictions upon the commanded satellite maneuver.
- the prescribed satellite maneuver profile 808 can be defined so as to limit the error between the predicted satellite attitude and the actual satellite attitude to less than 0.2 degrees and/or limit the yaw attitude rate and acceleration to a particular value.
- FIG. 9 An exemplary profile is illustrated in FIG. 9 as trace 902. Note that the difference between the ideal maneuver (trace 504) and the prescribed maneuver (trace 902) is small enough so that the solar panels 104 remain adequately directed at the Sun, yet by commanding the satellite attitude control system 200 to perform the prescribed maneuver instead of the ideal maneuver, the resulting actual satellite 100 attitude becomes predictable enough to allow the lever arm to be accurately predicted. A description of how the profile can be computed is discussed further below.
- spot beam gimbal angles are computed, as shown in block 816. This is accomplished using the predicted satellite 100 attitude over the time period T (computed in block 806), the gimbal geometry 810, and the target position of the spot beam antenna (e.g. the where the spot beam is aimed to direct it at one or more terrestrial targets) over the time period T, as computed in block 814.
- the position of the target spot beam over time period T is computed from a mission profile.
- the mission profile which describes the mission that one or more of the satellites 100 in the navigation system 700 is to perform (for example, which areas on the Earth's surface the spot beam antenna will be directed at), is preferably obtained by transmission from the ground segment 702 to the space segment 704, as illustrated in block 812.
- the gimbal geometry 810 (which includes, for example, a description of the dimensions and orientation of the gimbal structures and joints used to rotate the gimbals to direct the spot beam antenna 106S or equations describing the relationship between pointing direction and gimbal angles) can be loaded into a memory of the satellites 100 of the space segment 704, but can also be uplinked to the space segment 704 from the ground segment, if desired.
- the geometry can include, for example, positional vectors from the wide coverage antenna 106E phase center to the spot beam antenna the rotational axis of the first spot beam antenna 106S gimbal motor, and from this point to the rotational axis of the second spot beam antenna 106S gimbal motor, and to the second spot beam antenna 106S gimbal motor to the spot beam antenna 106 phase center.
- the spot beam gimbal angles are used to predict the lever arm 114 vector in a satellite bus-based coordinate system, as shown in block 820.
- the lever arm 114 vector L ⁇ A SV x t , y t , z t is transformed into Earth-centered Earth Fixed ECEF coordinates so that this information may be used by the navigational receivers 716 without further conversion. Since each spacecraft 100 is aware of its attitude, this operation is preferably performed on the spacecraft 100.
- the lever arm 114 vector is time dependent, the transmission of the lever arm correction at all times to the navigational receivers 716 would consume a significant amount of available transmission bandwidth.
- this difficulty is overcome by using curve-fitting techniques (including linear and non-linear regression) to identify closed form equations that describe the behavior of the lever arm over time during the time period T.
- the time-varying lever arm vector is expressed into its component parts (e.g. x ( t ), y ( t ), and z ( t )), and each time-varying component part is curve-fit to an equation definable by combination of functions, each defined by one or more functional coefficients. This is shown in block 824.
- the appropriate curve-fitting technique (e.g. the characteristics and order of the equation) can be selected according to the characteristics of the component part of the time varying lever arm 114 vector. For example, after z ( t ) for the time period T has been computed, z ( t ) can be evaluated to determine whether it is representable more accurately with a second, third, or fourth order polynomial. This can be accomplished, for example, by computing the mean squared error for each representation, and selecting the characteristics of the curve fit to minimize mean squared error. It may also be desirable to use different curve-fitting techniques and characteristics for the different component parts of the lever arm 114 vector.
- the lever arm vector coefficients are sent to the navigation receivers 716 in the user segment 706 via the spot beam from the spot beam antenna 106S, as shown in block 832.
- the navigation receivers receiving the coefficients 834 convert those coefficients to a time-varying lever arm correction prediction in ECEF coordinates, and apply the lever arm corrections at the appropriate time, as shown in blocks 834-838.
- the operations illustrated in FIGs. 8A and 8B as being performed by the space segment 704 can be implemented by software modules executed by the SCP 202, software modules executed by special or general purpose auxiliary processors in each satellite 100, or by dedicated electronic circuitry in each satellite 100.
- the operations shown in blocks 802, 806, 814, 816, 820, 822, and 824 are performed by a lever arm correction module 801, which can be implemented in the SCP 202, an auxiliary processor, or dedicated electronic circuitry.
- lever arm correction module 801 can include a satellite position predictor module 804, a satellite attitude predictor module 807, and a lever arm prediction module 826, which may further comprise a coordinate transformer 828 and curve fitting module 830 to perform the associated tasks shown in FIGs. 8A and 8B .
- These modules may also be implemented in the SCP 202, an auxiliary processor, or dedicated electronic circuitry.
- FIG. 10 is a diagram illustrating one technique that can be used to arrive at the prescribed maneuver profile.
- the performance of the satellite attitude control system is determined, using specifications and/or test data of the applicable components, including, for example, the reaction or momentum wheels, ACS thrusters, and attitude sensors.
- this data is used to determine the maximum angular rate and/or acceleration that can be achieved by the spacecraft 100 under the most demanding maneuvers it will be required to perform, typically the yaw maneuver at orbital noon and midnight for low Sun angles, as depicted in FIGs. 4A-4D . Since the yaw channel is typically most critical, a yaw steering profile can be examined. This can be accomplished using digital, analog, or hybrid control simulations of the satellite ACS 200.
- an ideal attitude profile is generated. This profile does not consider the limitations of the ACS, and yields zero pointing errors about all axes and zero pointing errors of the solar panels at all times.
- a candidate prescribed attitude profile is selected. The candidate profile is selected so as to attempt to match or closely approximate the ideal profile generated in block 1006.
- the candidate attitude profile is computed using closed form equations with parametric variability. The candidate profile will generally deviate from the ideal profile by an error, and the goal is to minimize this error.
- the candidate prescribed attitude profile is tested to determine if the spacecraft 100 can perform the maneuver described in the profile with sufficient (and typically nearly perfect) accuracy. Again, this can be performed with digital, analog, or hybrid simulations. If the test indicates that the satellite 100 can follow the prescribed maneuver profile, perhaps even better performance can be achieved by selecting a more demanding prescribed maneuver profile (using perhaps another candidate profile that follows the ideal profile more closely) that yields smaller errors when compared to the ideal profile, processing can return to block 1006 for the generation and eventual evaluation of that profile. If the test indicates that the satellite 100 cannot follow the prescribed maneuver profile, then the candidate prescribed maneuver profile is too demanding. The profile is adjusted in block 1014 and re-evaluated in steps 1008-1010. The foregoing process is iteratively repeated until a prescribed maneuver profile is found that achieves the least error between itself and the ideal maneuver profile, while still allowing the spacecraft to follow it essentially error free.
- ⁇ p a tan 2 ⁇ - sin ⁇ p , cos ⁇ p ⁇ sin ⁇
- ⁇ P the prescribed yaw attitude (angle)
- ⁇ the spacecraft orbit angle (i.e. its position)
- atan2 the two-argument arctangent function
- ⁇ p ⁇ a + ⁇ - ⁇ 2 ⁇ 2 n ⁇ sign ⁇ a ⁇ ⁇ min - ⁇ a
- ⁇ a is the actual current sun beta angle (i.e.
Landscapes
- Engineering & Computer Science (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
- Computer Networks & Wireless Communication (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Position Fixing By Use Of Radio Waves (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Description
- The present invention relates generally to space-based navigational systems, and in particular, to a space-based lever arm correction system applicable to navigational systems employing spot beams.
- The Global Positioning System (GPS) is a satellite system that transmits navigation signals that are received by ground-based GPS receivers and used to determine the position of the GPS receiver to a high degree of accuracy. GPS currently provides standard service to commercial receivers, and a higher accuracy service to military receivers authorized to receive such signals.
- In current systems, the GPS navigation signal is transmitted via a wide beam satellite antenna disposed on each GPS satellite. The wide beam antenna permits any GPS receiver having a line-of-sight to the GPS satellite to receive the navigational signal, and when the navigation signal from a sufficient number of GPS satellites has been acquired, the GPS receiver can determine its position via a precision clock and well-known triangulation techniques.
- Because GPS signals are also used in military applications, countermeasures can be expected to be applied in an attempt to reduce the effectiveness of the GPS system. One such countermeasure is jamming. To increase the effectiveness of the GPS signals in a jamming environment, a steerable high gain antenna may be used to transmit high intensity GPS signals via spot beams to areas where needed.
- One difficulty with this approach is that the high gain spot beam antenna is typically physically displaced from the wide beam antenna, and consequently, the phase center of each antenna is also displaced as well. This displacement is known as the "lever arm" between the antennas, and left uncorrected, can negatively affect the ability of the GPS receivers to determine their position. Without any correction, the lever arm between the wide beam antenna and a 7 meter diameter spot beam antenna can contribute up to 4.4 meters of user range error (URE). Depending on the GPS satellite constellation, this uncompensated URE can produce up to nine meters (RMS) of vertical (altitude) navigation error, which is a factor of 10 higher than the performance of the current GPS constellation of 0.9 meters (RMS). In civil aviation applications, such errors are sufficient to result in loss of life, and in military applications, they can result in increased collateral damage, and increased sortie and weapon consumption to perform the same mission.
- Further exacerbating this problem is the fact that in order to maintain proper Sun and Earth pointing, the GPS satellites are required to perform attitude maneuvers. Such maneuvers can be very large, particularly about the yaw axis.
- To achieve such high-accuracy navigation demanded for many missions, the GPS system must provide users with real-time corrections of the spot beam antenna phase center location relative to the Earth coverage antenna phase center, even while the spot beam antenna is moving to track specific terrestrial locations.
- What is needed is an apparatus and method for computing a correction for lever arm related errors, and for incorporating this correction in navigation computations. The present invention satisfies these needs.
- To address the requirements described above, the present invention discloses a method and apparatus for estimating a lever arm correction between a wide beam antenna and a spot beam antenna. The method comprises the steps of receiving a target spot beam position over the time period T in the satellite; generating a predicted position of the satellite over a time period T; generating predicted satellite attitude over the time period T from a prescribed satellite attitude maneuver profile over the time period T; generating predicted gimbal angles of the spot beam antenna over the time period T from the predicted satellite attitude and the target spot beam position over the time period T; predicting the lever arm correction between the wide beam antenna and the spot beam antenna over the time period T from the predicted gimbal angles of the spot beam antenna over the time period T and the predicted satellite attitude over the time period T; and transmitting the predicted lever arm correction from the satellite to a navigational receiver receiving the second navigation signal. The apparatus can be embodied in a navigational satellite, which comprises a first antenna transmitting a first navigational signal; a steerable second antenna transmitting a second navigational signal; a satellite receiver, for receiving data from a ground station, the data including a target spot beam position over a time period T; a lever arm correction module, for predicting a lever arm correction between the first antenna and the second antenna, and a navigation signal transmitter for transmitting the predicted lever arm correction to a navigational receiver.
- Referring now to the drawings in which like reference numbers represent corresponding parts throughout:
-
FIG. 1A is an illustration of a three-axis stabilized satellite; -
FIG. 1B is an illustration of an embodiment of a gimbal assembly; -
FIG. 2 is a diagram depicting a functional architecture of a representative satellite attitude control system; -
FIG. 3 is a diagram depicting geometrical relationships between the satellite, the Earth and the Sun; -
FIGs. 4A-4D are diagrams depicting the yaw noon maneuver at orbital midnight for a zero Sun angle; -
FIG. 5 is a diagram depicting the yaw attitude angle as a function of the orbit angle and the Sun angle; -
FIG. 6 is a diagram illustrating an exemplary embodiment of a technique for estimating the lever arm correction between the wide coverage antenna and the spot beam antenna; -
FIG. 7 is a diagram illustrating architectural elements of a space-based navigational system; -
FIGs. 8A and8B are diagrams illustrating exemplary process steps that can be used to implement one embodiment of the invention, and how these process steps can be allocated to the architectural elements; -
FIG. 9 is a plot of the yaw attitude angle of the satellite as a function of the orbit angle and the Sun angle for an exemplary prescribedsatellite 100 maneuver profile; and -
FIG. 10 is a diagram illustrating one technique that can be used to arrive at the prescribed maneuver profile. - In the following description, reference is made to the accompanying drawings which form a part hereof, and which is shown, by way of illustration, several embodiments of the present invention. It is understood that other embodiments may be utilized and structural changes may be made without departing from the scope of the present invention.
-
FIG. 1A illustrates a three-axis stabilized satellite orspacecraft 100. Thesatellite 100 has a main body 102 (which may be referred to as the "satellite bus"), one or moresolar panels 104, one or morenavigation beam antennas command antenna 108 which is used to communicate with a control ground station. Thesatellite 100 may also include one ormore sensors 110 to measure the attitude of thesatellite 100. These sensors may include sun sensors, earth sensors, and star sensors. Since the solar panels are often referred to by the designations "North" and "South", the solar panels inFIG. 1 are referred to by thenumerals - The three axes of the
spacecraft 100 are shown inFIG. 1A . The pitch axis lies along the line of the solar panels' 104N and 104S mutual rotation axes. The roll and yaw axes are perpendicular to the pitch axis and lie in the directions and planes shown. - In the illustrated embodiment, the
satellite 100 includes a firstnavigation beam antenna 106E and a secondnavigation beam antenna 106S. The firstnavigation beam antenna 106E is a wide-beam antenna which transmits a navigation signal with a beamwidth covering the widest range of the Earth's surface possible from that satellite's altitude at any time, and is directed toward the Earth along the yaw axis. Since thisantenna 106E offers the widest coverage of the Earth's surface, it is typically not steerable. The navigation system uses a constellation ofsuch satellites 100 to provide coverage anywhere on the Earth's surface by at least 4 different satellites at all times, thus permitting the navigational signal transmitted by the satellites to be used to determine the location of the receiver using triangulation techniques. - The second
navigation beam antenna 106S is a steerable spot beam antenna that provides a second navigation signal in a much narrower navigation beam. This allows transmission of a higher-strength beam to selected points on the ground without requiring excessive transmitter power, thus reducing the effectiveness of countermeasures such as jamming. Since the required service area includes substantially the entire surface of the Earth and the beamwidth of thespot beam antenna 106S is not wide enough to cover the entire surface area, the boresight of thespot beam antenna 106S can be steered about to direct the spot beam where desired. Such steering can be accomplished mechanically, by use of a gimbal structure driven by gimbal motors, or electronically, using phased arrays, for example. -
FIG. 1A also shows thephase center 112E of thewide beam antenna 106E and thephase center 112S of thespot beam antenna 106S. Since thespot beam antenna 106S is larger and offset from thewide beam antenna 106E, the phase centers 112S, 112E of the antennas are separated by anantenna lever arm 114, which is represented as a vector originating at thephase center 112E of thewide beam antenna 106E and extending to thephase center 112S of thespot beam antenna 106S. Due to motion of thesatellite bus 102 and thespot beam antenna 106S and other factors, theantenna lever arm 114 does not remain fixed, but can vary substantially over time. This variance is enough to add a significant uncertainty in the ability of a navigation receiver (such as a GPS receiver) to determine its location when a navigation signal is received via the spot beam. - The
spot beam antenna 106S may be steered electronically (by appropriate phasing of elements in a scanning array) or mechanically (by use of a non-scanning antenna and a gimbal assembly), or a combination of both.FIG. 1B illustrates an embodiment using a gimbal assembly 152 having aninner gimbal 154A and anouter gimbal 154B. Theinner gimbal 154A is associated with inner gimbal coordinateframe 156B while theouter gimbal 154B is associated with an outer gimbal coordinateframe 156C. Driven by gimbal motors or other devices (not illustrated), the inner andouter gimbals antenna 106S boresight (which is represented in antenna signal boresight coordinateframe 156D) where desired to transmit the navigation signal spot beam using theantenna 106S. Both inner and outer gimbals 154 typically include a potentiometer or other means to measure the gimbal angle. - To determine the angle at which the inner and outer gimbals 154 should be positioned to direct the
spot beam antenna 106S to the desired scan locations (to perform a specific mission profile), thespacecraft 100 determines its attitude via theattitude sensors 110, which may be mounted on the satellite bus orbody 102. Using the measured satellite attitude and the angular and translational displacement between thesatellite body 102 and thespot beam antenna 106S (as expressed by coordinatesystems 156A-156D), thesatellite 100 can determine the appropriate gimbal angles, and commands the gimbal motors to move the gimbals 156 to the appropriate positions. -
FIG. 2 is a diagram depicting the functional architecture of a representativeattitude control system 200. Control of the spacecraft is provided by a computer or spacecraft control processor (SCP) 202. TheSCP 202 performs a number of functions which may include post ejection sequencing, transfer orbit processing, acquisition control, station-keeping control, normal mode control, mechanisms control, fault protection, and spacecraft systems support, among others. - The
SCP 202 may implement one or more processing modules such asantenna control module 276, which is used to control the satellite spotbeam antenna drive 274 to slew thespot beam antenna 106S to the appropriate orientation and to transmit a navigation signal. Alternatively, theantenna control module 276 can be implemented in a different processor or in dedicated circuitry. - Input to the
spacecraft control processor 202 may come from any combination of a number of spacecraft components and subsystems, such as a transferorbit sun sensor 204, anacquisition sun sensor 206, aninertial reference unit 208, a transferorbit Earth sensor 210, an operationalorbit Earth sensor 212, a normal mode wideangle sun sensor 214, amagnetometer 216, and one ormore star sensors 218. - The
SCP 202 generates control signal commands 220 which are directed to acommand decoder unit 222. The command decoder unit operates the load shedding andbattery charging systems 224. The command decoder unit also sends signals to the magnetic torque control unit (MTCU) 226 and thetorque coil 228. - Wheel torque commands 262 are generated by the
SCP 202 and are communicated to thewheel drive electronics 240 which command the speed of the reaction wheels in reaction wheel assembly(s) 244. Typically, thespacecraft 100 includes four reaction wheels, at least one in each orthogonal direction, and one for redundancy purposes. The speed of the reaction wheels is also measured and fed back to theSCP 202 byfeedback control signal 264. TheSCP 202 also communicates commands anddata 254 with command stations (further described in connection withFIG. 7 ) via a satellite transmitter/receiver (or transceiver) 258. - The foregoing describes an exemplary space stabilized satellite attitude control system. The present invention can be implemented with other attitude control system designs as well.
- The
SCP 202 may include or have access tomemory 270, such as a random access memory (RAM). Generally, theSCP 202 operates under control of anoperating system 272 stored in thememory 270, and interfaces with the other system components to accept inputs and generate outputs, including commands. Applications running in theSCP 202 access and manipulate data stored in thememory 270. Thespacecraft 100 may also comprise an external communication device such as a satellite link for communicating with other computers at, for example, a ground station. If necessary, operation instructions for new applications can be uploaded from ground stations. - In one embodiment, instructions implementing the
operating system 272, application programs, and other modules are tangibly embodied in a computer-readable medium, e.g., data storage device, which could include a RAM, EEPROM, or other memory device. Further, theoperating system 272 and the computer program are comprised of instructions which, when read and executed by theSCP 202, causes thespacecraft processor 202 to perform the steps necessary to implement and/or use the present invention. Computer program and/or operating instructions may also be tangibly embodied inmemory 270 and/or data communications devices (e.g. other devices in thespacecraft 100 or on the ground), thereby making a computer program product or article of manufacture according to the invention. As such, the terms "program storage device," "article of manufacture" and "computer program product" as used herein are intended to encompass a computer program accessible from any computer readable device or media. - As described above, it is desirable to correct for navigation errors caused by the
lever arm 114 between the wide beam antenna and the steerable spot beam antenna. One approach taken to correct for such errors is to compute theantenna lever arm 114 in the satellite coordinate frame, and transmit this information to the GPS receivers in the navigational message. The GPS receiver then uses this information to compute corrections in the GPS (e.g. Earth Centered Earth Fixed) coordinate frame. However, this technique is complicated by the maneuvers that thesatellite 100 must undergo to direct thesolar panels 104 at the Sun and thewide beam antenna 106E at the Earth. Depending on the Sun angle, this requires a rather substantial yaw maneuver twice each orbit (at orbit noon and midnight). -
FIG. 3 is a diagram depicting the geometrical relationships between thesatellite 100, theEarth 308 and theSun 306. Eachsatellite 100 of the GPS system is in anorbit 302 around theEarth 308, thus defining anorbital plane 304. The angle θ between theorbital plane 304 and theSun 306 is the Sun angle. Thesatellite 100 orbits theEarth 308, always directing theantenna 106E at theEarth 308. Thesatellite 100 attitude andsolar panels 104 are adjusted to keep thesolar panels 104 directed at theSun 306 and theantenna 106E directed at theEarth 308 at the same time. Typically, this requires thesatellite 100 to perform a large rotation about the yaw axis at the orbital midnight (0 degree orbit angle) and the orbital noon (180 degree orbit angle). The smaller the Sun angle θ, the larger the required rotation about the yaw axis. -
FIGs. 4A-4D are diagrams depicting the yaw turn maneuver at orbital midnight for a zero Sun angle (wherein theSun 306 is in the orbital plane 304).FIG. 4A illustrates thesatellite 100 attitude andsolar panel 104 orientation as the orbital midnight is approached. Note that thesolar panels 104 are canted at an angle to direct them to be perpendicular to the Sun's rays. Shortly before theSun 306 begins to be eclipsed by theEarth 308, thesatellite 100 begins its maneuver and yaws counter clockwise as shown inFIG. 4B through FIG. 4D . When theSun 306 fully emerges from the eclipse, thesatellite 100 has yawed 180 degrees from its initial position. -
FIG. 5 is a diagram plotting the ideal yaw angle as a function of the orbit angle. Athigher Sun 306 angles, the yaw angle movement is gradual, as shown intrace 502. However, atlow Sun 306 angles, the yaw angle abruptly transitions 180 degrees at orbital angles of 0 and 180 degrees, with resulting high yaw angular rates and/or accelerations, as shown intrace 504. Such low Sun angle conditions (e.g. less than about five degrees) occur over a significant portion of the year. For example, the eclipse season occurs twice a year for each orbital plane. Since a GPS constellation typically includes six orbital planes, the constellation typically experiences eclipse conditions in one of its planes at all times of the year. -
FIG. 6 is a diagram illustrating an exemplary embodiment of a technique for estimating thelever arm 114 correction between thewide coverage antenna 106E and thespot beam antenna 106S. Inblock 602, a prediction of thesatellite 100 position over a time period T is generated.Block 604 illustrates the generation of a prediction of thesatellite 100 attitude over the time period T from a profile of a prescribed navigational satellite attitude maneuver. Finally, block 606 illustrates predicting the lever arm correction between the wide coverage antenna and the spot beam antenna over the time period T using the predicted satellite attitude over the time period T and a target position on the Earth of the spot beam signal. -
FIG. 7 is a diagram illustrating architectural elements of a space-basednavigational system 700. Thenavigational system 700 comprises aspace segment 704 having a plurality of navigational satellites 100 (hereinafter alternatively referred to as spacecraft) in orbit around the Earth. Each of thesatellites 100 includes afirst transmitter 710, which transmits a first navigational signal using thewide beam antenna 106E and asecond transmitter 712, which, using thespot beam antenna 106S, transmits aspot beam 719 having a second navigational signal to one or more of thenavigational receivers 716 that together form theuser segment 706. Thespot beam 719 covers an area 715 on the Earth's surface, and is centered on the targetspot beam position 717. Both the first navigational signal and the second navigational signal include information that can be used to determine the location of thenavigational receiver 716. In a jamming environment, a navigation receiver may lose lock on the first navigation signal and will not be able to produce any positional solution. The second navigational signal, however, is typically of greater power by virtue of being transmitted via a spot beam, and, will be used for positional determination using thenavigational receiver 716 in the jamming environment. Thesatellites 100 in thespace segment 704 are controlled via commands from theground segment 702, which comprises one or more ground-basedcommand stations 707, each including acommand center 708 and a communicatively coupledantenna 714 for transmitting and receiving information and commands to and from thesatellites 100 of thespace segment 704. - The operations illustrated in
FIG. 6 can be performed on-board thesatellite 100, theground segment 702 controlling thesatellite 100, by thenavigational receivers 716 making up theuser segment 706, or any combination thereof. - However, while a fixed
antenna lever arm 114 can be computed in the satellite coordinate frame 156 and transmitted to the GPS receiver in the navigational message assuming an idealized yaw steering profile, this technique has serious disadvantages. First, this approach also cannot be applied to GPS systems that use steerable spot beam antennas, because the GPS receivers do not know thesatellite 100 attitude, very accurately, and thus have insufficient information to compute an accurate estimate of thelever arm 114 at all times. Second, even if the GPS receivers were provided with thesatellite 100 attitude, this technique requires that the GPS receivers perform the majority of the correction processing. Since the GPS receivers are often required to be very small, lightweight, and low power consumption devices, such processing can be a significant problem. Also, because of satelliteattitude control system 200 limitations, the high yaw rate maneuvers that occur atlow Sun 306 angles (such as shown inFIG. 5 ) add substantial uncertainty to theantenna lever arm 114 corrections transmitted to the GPS receivers. - One solution to these problems is to compute
discrete lever arm 114 corrections in the GPS coordinate frame in thesatellite 100 and periodically transmit the corrections to the GPS receivers. Since the correction data is in the GPS coordinate frame, it can be applied to the satellite orbital motion data already provided by the GPS system, and thus requires only minimal GPS receiver processing. However, such discrete corrections do not provide solutions in between the discrete points, and adding discrete points would consume a prohibitively large portion of the transmission bandwidth in the navigation message. - Finally, a problem with either of the foregoing correction approaches is that the corrections must be extrapolated for some period into the future. The first approach (that of transmitting the fixed antenna lever arm position in the
satellite 100 coordinate frame in the navigational message and relying upon the GPS receivers to use this information along with a satellite yaw steering maneuver to enable computation of the corrections in a GPS coordinate frame), relies on an idealized (and often inaccurate) prediction of thesatellite 100 attitude for all relevant times. The second approach makes predictions at least 15 minutes into the future, and during those 15 minutes, without additional information or constraints, this solution can yield large errors. -
FIGs. 8A and8B are diagrams illustrating exemplary process steps that can be used to implement one embodiment of the invention, and how these process steps can be allocated to the architectural elements (e.g. thecontrol segment 702, thespace segment 704, and the user segment 706). - Referring first to
FIG. 8A , the predicted position of thesatellite 100 over a time period T is computed, as shown inblock 802. The predicted position of thesatellite 100 can be determined using a variety of methods well known in the art, and can include parameters to model satellite maneuvers (whether for mission specific or station keeping purposes). Also, for a navigational satellite such as a GPS satellite, the parameters for predicting satellite position are normally present on-board the satellite as an integral part of the navigational system. The predicted position of thesatellite 100 can be determined from this information. - Next, as shown in
block 806, a prediction of the satellite's attitude over the same time period T is computed. This is accomplished using the computed satellite position (from the computation shown in block 802), and from a prescribedsatellite maneuver profile 808 that can be obtained either from theground segment 702 or the memory of thesatellite 100 in thespace segment 704. -
FIG. 9 is a plot of the yaw attitude angle of thesatellite 100 as a function of theorbit angle 310 and theSun 306 angle for an exemplaryprescribed satellite 100 maneuver profile. The prescribed maneuver profile can be defined such that, when accounting for the capabilities and limitations of the satellite'sattitude control system 200, thesatellite 100 can execute the prescribed attitude maneuver within a specific angular tolerance, the selection of such tolerance is a function of the desired lever arm estimation accuracy. In one embodiment, the value of this tolerance, expressed as the difference between the predicted satellite attitude and the actual satellite attitude to less than 0.2 degrees. The prescribed maneuver profile can be defined so as to limit the maximum angular velocity and/or acceleration about all axes (pitch, roll, and yaw) not to exceed maximum values or such that the maximum angular velocity and/or acceleration does not exceed maximum values in any one of the axes (for example, in the yaw axis). - As was true in
FIG. 5 ,trace 502 illustrates yaw angle profile for ahigh Sun 306 angle θ. Athigh Sun 306 angles θ, the required yaw rates and accelerations are relatively small. However, atlow Sun 306 angles θ, the required yaw rates and accelerations are much higher, as shown intrace 504. These yaw rates are high enough so that the satellite'sattitude control system 200 cannot follow the ideal profile, and thus, the deviation from the expectedsatellite 100 attitude angle and theactual satellite 100 attitude angle is significant enough to negatively impact the estimation of thelever arm 114. Also, the maneuver is initiated based upon sensor (e.g. Sun, Earth, and/or star sensors) measurements, hence the time the maneuver is initiated is another source of unpredictability. To prevent this, the satellite is commanded to perform aprescribed satellite maneuver 808. - This prescribed
satellite maneuver profile 808 can be defined such that, when accounting for the capabilities and limitation of the satellite'sattitude control system 200, thesatellite 100 can execute the prescribed attitude maneuver to within specified tolerances. The maneuver profile can be arrived at by (1) limiting the satellite attitude angular rate and/or acceleration in all axes (e.g. pitch, roll, and yaw) or a single axis (e.g. yaw only) to a maximum value, (2) limiting the error between the predicted satellite attitude and the actual satellite attitude (whenattitude control system 200 limitations are considered), (3) and/or placing other suitable restrictions upon the commanded satellite maneuver. For example, the prescribedsatellite maneuver profile 808 can be defined so as to limit the error between the predicted satellite attitude and the actual satellite attitude to less than 0.2 degrees and/or limit the yaw attitude rate and acceleration to a particular value. - An exemplary profile is illustrated in
FIG. 9 astrace 902. Note that the difference between the ideal maneuver (trace 504) and the prescribed maneuver (trace 902) is small enough so that thesolar panels 104 remain adequately directed at the Sun, yet by commanding the satelliteattitude control system 200 to perform the prescribed maneuver instead of the ideal maneuver, the resultingactual satellite 100 attitude becomes predictable enough to allow the lever arm to be accurately predicted. A description of how the profile can be computed is discussed further below. - Returning to
FIG. 8A , spot beam gimbal angles are computed, as shown inblock 816. This is accomplished using the predictedsatellite 100 attitude over the time period T (computed in block 806), thegimbal geometry 810, and the target position of the spot beam antenna (e.g. the where the spot beam is aimed to direct it at one or more terrestrial targets) over the time period T, as computed inblock 814. The position of the target spot beam over time period T is computed from a mission profile. The mission profile, which describes the mission that one or more of thesatellites 100 in thenavigation system 700 is to perform (for example, which areas on the Earth's surface the spot beam antenna will be directed at), is preferably obtained by transmission from theground segment 702 to thespace segment 704, as illustrated inblock 812. - The gimbal geometry 810 (which includes, for example, a description of the dimensions and orientation of the gimbal structures and joints used to rotate the gimbals to direct the
spot beam antenna 106S or equations describing the relationship between pointing direction and gimbal angles) can be loaded into a memory of thesatellites 100 of thespace segment 704, but can also be uplinked to thespace segment 704 from the ground segment, if desired. The geometry can include, for example, positional vectors from thewide coverage antenna 106E phase center to the spot beam antenna the rotational axis of the firstspot beam antenna 106S gimbal motor, and from this point to the rotational axis of the secondspot beam antenna 106S gimbal motor, and to the secondspot beam antenna 106S gimbal motor to the spot beam antenna 106 phase center. - Turning to
FIG. 8B , the spot beam gimbal angles are used to predict thelever arm 114 vector in a satellite bus-based coordinate system, as shown inblock 820. Inblock 828, thelever arm 114 vectornavigational receivers 716 without further conversion. Since eachspacecraft 100 is aware of its attitude, this operation is preferably performed on thespacecraft 100. - Because the
lever arm 114 vector is time dependent, the transmission of the lever arm correction at all times to thenavigational receivers 716 would consume a significant amount of available transmission bandwidth. In a preferred embodiment of the invention, this difficulty is overcome by using curve-fitting techniques (including linear and non-linear regression) to identify closed form equations that describe the behavior of the lever arm over time during the time period T. In particular, the time-varying lever arm vector is expressed into its component parts (e.g. x(t), y(t), and z(t)), and each time-varying component part is curve-fit to an equation definable by combination of functions, each defined by one or more functional coefficients. This is shown inblock 824. -
- The appropriate curve-fitting technique (e.g. the characteristics and order of the equation) can be selected according to the characteristics of the component part of the time varying
lever arm 114 vector. For example, after z(t) for the time period T has been computed, z(t) can be evaluated to determine whether it is representable more accurately with a second, third, or fourth order polynomial. This can be accomplished, for example, by computing the mean squared error for each representation, and selecting the characteristics of the curve fit to minimize mean squared error. It may also be desirable to use different curve-fitting techniques and characteristics for the different component parts of thelever arm 114 vector. - Next, the lever arm vector coefficients are sent to the
navigation receivers 716 in theuser segment 706 via the spot beam from thespot beam antenna 106S, as shown inblock 832. The navigation receivers receiving thecoefficients 834, convert those coefficients to a time-varying lever arm correction prediction in ECEF coordinates, and apply the lever arm corrections at the appropriate time, as shown in blocks 834-838. - The operations illustrated in
FIGs. 8A and8B as being performed by thespace segment 704 can be implemented by software modules executed by theSCP 202, software modules executed by special or general purpose auxiliary processors in eachsatellite 100, or by dedicated electronic circuitry in eachsatellite 100. For example, in one embodiment, the operations shown inblocks arm correction module 801, which can be implemented in theSCP 202, an auxiliary processor, or dedicated electronic circuitry. Further, the leverarm correction module 801 can include a satelliteposition predictor module 804, a satelliteattitude predictor module 807, and a leverarm prediction module 826, which may further comprise a coordinatetransformer 828 and curvefitting module 830 to perform the associated tasks shown inFIGs. 8A and8B . These modules may also be implemented in theSCP 202, an auxiliary processor, or dedicated electronic circuitry. -
FIG. 10 is a diagram illustrating one technique that can be used to arrive at the prescribed maneuver profile. Inblock 1002, the performance of the satellite attitude control system (ACS) is determined, using specifications and/or test data of the applicable components, including, for example, the reaction or momentum wheels, ACS thrusters, and attitude sensors. Inblock 1004, this data is used to determine the maximum angular rate and/or acceleration that can be achieved by thespacecraft 100 under the most demanding maneuvers it will be required to perform, typically the yaw maneuver at orbital noon and midnight for low Sun angles, as depicted inFIGs. 4A-4D . Since the yaw channel is typically most critical, a yaw steering profile can be examined. This can be accomplished using digital, analog, or hybrid control simulations of thesatellite ACS 200. Inblock 1006, an ideal attitude profile is generated. This profile does not consider the limitations of the ACS, and yields zero pointing errors about all axes and zero pointing errors of the solar panels at all times. Inblock 1008, a candidate prescribed attitude profile is selected. The candidate profile is selected so as to attempt to match or closely approximate the ideal profile generated inblock 1006. Preferably, the candidate attitude profile is computed using closed form equations with parametric variability. The candidate profile will generally deviate from the ideal profile by an error, and the goal is to minimize this error. - Using this information, the candidate prescribed attitude profile is tested to determine if the
spacecraft 100 can perform the maneuver described in the profile with sufficient (and typically nearly perfect) accuracy. Again, this can be performed with digital, analog, or hybrid simulations. If the test indicates that thesatellite 100 can follow the prescribed maneuver profile, perhaps even better performance can be achieved by selecting a more demanding prescribed maneuver profile (using perhaps another candidate profile that follows the ideal profile more closely) that yields smaller errors when compared to the ideal profile, processing can return to block 1006 for the generation and eventual evaluation of that profile. If the test indicates that thesatellite 100 cannot follow the prescribed maneuver profile, then the candidate prescribed maneuver profile is too demanding. The profile is adjusted inblock 1014 and re-evaluated in steps 1008-1010. The foregoing process is iteratively repeated until a prescribed maneuver profile is found that achieves the least error between itself and the ideal maneuver profile, while still allowing the spacecraft to follow it essentially error free. - An example candidate prescribed yaw attitude profile is given below
where ψP is the prescribed yaw attitude (angle), α is the spacecraft orbit angle (i.e. its position), atan2 is the two-argument arctangent function and
where βa is the actual current sun beta angle (i.e. the angle between the orbit plane and the sun), βmin is a minimum sun beta angle below which the satellite cannot perfectly follow the ideal profile, and the orbit angle α is normalized ("unwrapped") so that its value always falls between -π and +π. The exponent n is the parameter of the above equations, and it would be adjusted on each pass of the blocks 1008-1012 to achieve the best performance. Note that for this example there are some constraints on n, such as it must be even and positive. While other forms of the candidate profile may yield better performance, the above technique and candidate profile yields a value of n = 10. - This concludes the description of the preferred embodiments of the present invention. The foregoing description of the preferred embodiment of the invention has been presented for the purposes of illustration and description. It is not intended to be exhaustive or to limit the invention to the precise form disclosed. Many modifications and variations are possible in light of the above teaching. It is intended that the scope of the invention be limited not by this detailed description, but rather by the claims appended hereto.
Claims (10)
- A method of estimating a lever arm (114) correction between a first wide beam antenna (106E) and a steerable second spot beam antenna (106S) of a navigational satellite (100), the first antenna (106E) transmitting a first navigational signal and the steerable second antenna (106S) transmitting a second navigational signal, comprising the steps of:receiving from a ground station a target position of the spot beam (717) over the time period T in the satellite (100);generating a predicted position of the satellite (100) over a time period T;generating predicted satellite attitude over the time period T from a profile (902) of a prescribed satellite maneuver over the time period T;generating predicted gimbal angles of the spot beam antenna (106S) over the time period T from the predicted satellite attitude and the target position of the spot beam (717) over the time period T;predicting the lever arm (114) correction between the wide beam antenna (106E) and the spot beam antenna (106S) over the time period T from the predicted gimbal angles of the spot beam antenna over the time period T and the predicted satellite attitude over the time period T; andtransmitting the predicted lever arm correction from the satellite (100) to a navigational receiver (716).
- The method of claim 1, wherein the predicted satellite attitude is generated from an angle between an orbital plane (304) of the at least one satellite (100) and the Sun (306) and the predicted position of the at least one satellite (100) within the orbital plane (304).
- The method of claim 1, wherein the prescribed satellite maneuver is constrained to limit an attitude angular rate of the navigational satellite (100) to a maximum value.
- The method of claim 1, wherein the prescribed satellite maneuver constrained to limit an attitude angular acceleration of the navigational satellite (100) to a maximum value.
- The method of claim 1, wherein the prescribed satellite maneuver is constrained such that when the at least one navigational satellite (100) executes the prescribed maneuver, the at least one navigational satellite (100) can concurrently direct a satellite solar panel (104) to the Sun (306) and the wide beam antenna (106E) to the Earth (308).
- A navigational satellite (100), comprising:a first wide beam antenna (106E) transmitting a first navigational signal;a steerable second spot beam antenna (106S) transmitting a second navigational signal;a satellite receiver (716), for receiving data from a ground station (707), the data including a target position of the spot beam (717) over a time period T;a lever arm correction module (801), for predicting a lever arm correction between the first antenna (106E) and the second antenna (106S), the lever arm correction module (801) comprising:a satellite position predictor module (804), for generating a predicted satellite position over the time period T;a satellite attitude predictor module (807), for generating a predicted satellite attitude over the time period T from a profile (902) of a prescribed satellite maneuver over the time period T;a lever arm prediction module (826), for predicting the lever arm correction between the first antenna and the second antenna over the time period T from the target position of the spot beam over the time period T and the satellite attitude over the time period T;a transmitter (710, 712) for transmitting the predicted lever arm correction to a navigational receiver.
- The navigational satellite (100) of claim 6, wherein the prescribed maneuver is constrained such that when the at least one navigational satellite (100) follows the prescribed maneuver, the at least one navigational satellite (100) can concurrently direct a satellite solar panel (104) to the Sun (306) and the wide beam antenna (106E) to the Earth (308).
- The navigational satellite (100) of claim 6, wherein the predicted gimbal angles of the spot beam antenna (106S) over time T are further generated from spot beam antenna geometry.
- The navigational satellite (100) of claim 6, wherein the target spot beam position (717) changes during the time period T.
- The navigational satellite (100) of claim 6, wherein the lever arm prediction module (826) computes the time-varying lever arm correction vector in navigational satellite coordinates, and the lever arm prediction module (826) further comprises a curve fitting module (830) for curve fitting the time-varying lever arm correction vector to produce lever arm correction vector coefficients.
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US11/247,641 US7221317B2 (en) | 2005-10-10 | 2005-10-10 | Space-based lever arm correction in navigational systems employing spot beams |
Publications (2)
Publication Number | Publication Date |
---|---|
EP1772742A1 EP1772742A1 (en) | 2007-04-11 |
EP1772742B1 true EP1772742B1 (en) | 2012-05-16 |
Family
ID=37451221
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP06255096A Not-in-force EP1772742B1 (en) | 2005-10-10 | 2006-10-02 | Correction of the distance between phase centres of two directional antenneas of a navigational satellite |
Country Status (2)
Country | Link |
---|---|
US (1) | US7221317B2 (en) |
EP (1) | EP1772742B1 (en) |
Families Citing this family (12)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8713324B2 (en) * | 2006-01-18 | 2014-04-29 | Overhorizon (Cyprus) Plc | Systems and methods for tracking mobile terrestrial terminals for satellite communications |
US7667645B2 (en) | 2006-05-25 | 2010-02-23 | The Boeing Company | GPS gyro calibration |
US7423585B2 (en) * | 2006-05-25 | 2008-09-09 | The Boeing Company | Navigation signal group delay calibration |
US7925439B2 (en) * | 2006-10-19 | 2011-04-12 | Topcon Positioning Systems, Inc. | Gimbaled satellite positioning system antenna |
US7847588B2 (en) * | 2008-08-14 | 2010-12-07 | Nantero, Inc. | Nonvolatile nanotube programmable logic devices and a nonvolatile nanotube field programmable gate array using same |
US20100088063A1 (en) * | 2008-10-01 | 2010-04-08 | A-Tech Corporation | Method and Apparatus for Precision Azimuth Measurement |
FR2962811B1 (en) * | 2010-07-19 | 2013-09-27 | Thales Sa | SYNTHETIC OPENING ANTENNA DEVICE FOR SIGNAL TRANSMISSION OF A SATELLITE NAVIGATION SYSTEM COMPRISING A CARRIER AND MEANS FOR DETERMINING ITS PATHWAY |
FR2997386B1 (en) * | 2012-10-31 | 2015-05-29 | Thales Sa | OPTIMIZED PROPULSION DEVICE FOR ORBIT CONTROL AND SATELLITE ATTITUDE CONTROL |
US9376221B1 (en) | 2012-10-31 | 2016-06-28 | The Boeing Company | Methods and apparatus to point a payload at a target |
US20150355340A1 (en) * | 2013-01-21 | 2015-12-10 | Leica Geosystems Ag | Co-located antenna |
CN110579736B (en) * | 2018-06-08 | 2021-10-08 | 广州极飞科技股份有限公司 | Method and device for correcting direction of machine head and flight controller |
CN117439656B (en) * | 2023-12-08 | 2024-02-23 | 成都时代宇辰科技有限公司 | Automatic correction method for absolute coordinate system of VICTS satellite communication antenna |
Family Cites Families (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5184139A (en) * | 1990-08-29 | 1993-02-02 | Kabushiki Kaisha Toshiba | Antenna pointing equipment |
US5928309A (en) * | 1996-02-05 | 1999-07-27 | Korver; Kelvin | Navigation/guidance system for a land-based vehicle |
-
2005
- 2005-10-10 US US11/247,641 patent/US7221317B2/en not_active Expired - Fee Related
-
2006
- 2006-10-02 EP EP06255096A patent/EP1772742B1/en not_active Not-in-force
Also Published As
Publication number | Publication date |
---|---|
EP1772742A1 (en) | 2007-04-11 |
US7221317B2 (en) | 2007-05-22 |
US20070080859A1 (en) | 2007-04-12 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
EP1772742B1 (en) | Correction of the distance between phase centres of two directional antenneas of a navigational satellite | |
US7221316B2 (en) | Control segment-based lever-arm correction via curve fitting for high accuracy navigation | |
US7095376B1 (en) | System and method for pointing and control of an antenna | |
US10483629B1 (en) | Antenna beam pointing system | |
EP0683098B1 (en) | Spacecraft attitude determination using sun sensor, earth sensor and space-to-ground link | |
US8213803B2 (en) | Method and system for laser based communication | |
US7357356B1 (en) | Attitude and antenna steering system for geosynchronous earth orbit (GEO) spacecraft | |
US7877173B2 (en) | Method and apparatus for determining a satellite attitude using crosslink reference signals | |
US6691033B1 (en) | System and method for calibrating inter-star-tracker misalignments in a stellar inertial attitude determination system | |
US7062363B2 (en) | Refinement of spacecraft angular velocity and attitude estimates using star data | |
US7256734B2 (en) | Spot beam antenna boresight calibration using GPS receivers | |
US6615117B2 (en) | Attitude determination system and method with outer-loop gyro scale-factor non-linearity calibration | |
EP0544198A1 (en) | Method and apparatus for controlling a solar wing of a satellite using a sun sensor | |
US6504502B1 (en) | Method and apparatus for spacecraft antenna beam pointing correction | |
US6771217B1 (en) | Phased array pointing determination using inverse pseudo-beacon | |
US7258306B2 (en) | Thermal deformation determination for payload pointing using space-based beacon | |
US6441776B1 (en) | Method and apparatus for spacecraft payload pointing registration | |
US6567040B1 (en) | Offset pointing in de-yawed phased-array spacecraft antenna | |
US7124001B2 (en) | Relative attitude estimator for multi-payload attitude determination | |
US6216983B1 (en) | Ephemeris/attitude reference determination using communications links | |
KR20010039649A (en) | Ephemeris/attitude reference determination using on-board optics and other satellite ephemeris | |
US6762716B2 (en) | Digital beacon asymmetry and quantization compensation | |
US7129889B1 (en) | User segment-based lever arm correction via prescribed maneuver for high-accuracy navigation | |
US6566640B2 (en) | System and method for autonomous correction of star tracker high spatial frequency (HSF) error | |
JP4295618B2 (en) | Satellite attitude adjustment for GPS initialization |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
PUAI | Public reference made under article 153(3) epc to a published international application that has entered the european phase |
Free format text: ORIGINAL CODE: 0009012 |
|
17P | Request for examination filed |
Effective date: 20061007 |
|
AK | Designated contracting states |
Kind code of ref document: A1 Designated state(s): AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HU IE IS IT LI LT LU LV MC NL PL PT RO SE SI SK TR |
|
AX | Request for extension of the european patent |
Extension state: AL BA HR MK YU |
|
AKX | Designation fees paid |
Designated state(s): AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HU IE IS IT LI LT LU LV MC NL PL PT RO SE SI SK TR |
|
17Q | First examination report despatched |
Effective date: 20071129 |
|
GRAP | Despatch of communication of intention to grant a patent |
Free format text: ORIGINAL CODE: EPIDOSNIGR1 |
|
RIC1 | Information provided on ipc code assigned before grant |
Ipc: G01S 1/00 20060101AFI20111109BHEP Ipc: G01S 19/02 20100101ALI20111109BHEP Ipc: H01Q 1/12 20060101ALI20111109BHEP Ipc: H01Q 3/00 20060101ALI20111109BHEP |
|
GRAS | Grant fee paid |
Free format text: ORIGINAL CODE: EPIDOSNIGR3 |
|
GRAA | (expected) grant |
Free format text: ORIGINAL CODE: 0009210 |
|
AK | Designated contracting states |
Kind code of ref document: B1 Designated state(s): AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HU IE IS IT LI LT LU LV MC NL PL PT RO SE SI SK TR |
|
REG | Reference to a national code |
Ref country code: GB Ref legal event code: FG4D |
|
REG | Reference to a national code |
Ref country code: CH Ref legal event code: EP |
|
REG | Reference to a national code |
Ref country code: AT Ref legal event code: REF Ref document number: 558334 Country of ref document: AT Kind code of ref document: T Effective date: 20120615 |
|
REG | Reference to a national code |
Ref country code: IE Ref legal event code: FG4D |
|
REG | Reference to a national code |
Ref country code: DE Ref legal event code: R096 Ref document number: 602006029409 Country of ref document: DE Effective date: 20120712 |
|
REG | Reference to a national code |
Ref country code: NL Ref legal event code: VDEP Effective date: 20120516 |
|
REG | Reference to a national code |
Ref country code: LT Ref legal event code: MG4D Effective date: 20120516 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: CY Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 Ref country code: LT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 Ref country code: SE Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 Ref country code: IS Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120916 Ref country code: PL Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 Ref country code: FI Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 |
|
REG | Reference to a national code |
Ref country code: AT Ref legal event code: MK05 Ref document number: 558334 Country of ref document: AT Kind code of ref document: T Effective date: 20120516 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: PT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120917 Ref country code: SI Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 Ref country code: LV Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 Ref country code: GR Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120817 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: BE Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: NL Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 Ref country code: CZ Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 Ref country code: DK Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 Ref country code: EE Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 Ref country code: AT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 Ref country code: RO Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 Ref country code: SK Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: IT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 |
|
PLBE | No opposition filed within time limit |
Free format text: ORIGINAL CODE: 0009261 |
|
STAA | Information on the status of an ep patent application or granted ep patent |
Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT |
|
26N | No opposition filed |
Effective date: 20130219 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: ES Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120827 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: MC Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20121031 |
|
REG | Reference to a national code |
Ref country code: CH Ref legal event code: PL |
|
REG | Reference to a national code |
Ref country code: DE Ref legal event code: R097 Ref document number: 602006029409 Country of ref document: DE Effective date: 20130219 |
|
REG | Reference to a national code |
Ref country code: IE Ref legal event code: MM4A |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: BG Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120816 Ref country code: LI Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20121031 Ref country code: IE Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20121002 Ref country code: CH Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20121031 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: TR Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20120516 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: LU Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20121002 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: HU Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20061002 |
|
REG | Reference to a national code |
Ref country code: FR Ref legal event code: PLFP Year of fee payment: 10 |
|
REG | Reference to a national code |
Ref country code: FR Ref legal event code: PLFP Year of fee payment: 11 |
|
REG | Reference to a national code |
Ref country code: FR Ref legal event code: PLFP Year of fee payment: 12 |
|
PGFP | Annual fee paid to national office [announced via postgrant information from national office to epo] |
Ref country code: FR Payment date: 20171025 Year of fee payment: 12 Ref country code: DE Payment date: 20171027 Year of fee payment: 12 |
|
PGFP | Annual fee paid to national office [announced via postgrant information from national office to epo] |
Ref country code: GB Payment date: 20171027 Year of fee payment: 12 |
|
REG | Reference to a national code |
Ref country code: DE Ref legal event code: R119 Ref document number: 602006029409 Country of ref document: DE |
|
GBPC | Gb: european patent ceased through non-payment of renewal fee |
Effective date: 20181002 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: DE Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20190501 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: FR Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20181031 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: GB Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20181002 |